It is common knowledge that the aeronautical industry requires structures that on the one hand withstand the loads to which they are subjected, meeting high requirements on strength and stiffness, and on the other hand are as light as possible. A consequence of this requirement is the more and more extensive use of composite materials in primary structures, because with appropriate application of said composite materials it is possible in consequence to achieve an important weight saving relative to a design in metallic material.
Integrated structures in particular have proved to be very efficient in this respect. We speak of an integrated structure when the different structural elements are manufactured in one step. This is another advantage of the use of composite materials because with their condition of independent layers that can be lay-up in the desired form, they offer the possibility of increasing integration of the structure, which moreover often produces a cost saving—equally essential when competing in the marketplace—as there are fewer individual parts to be assembled.
The main structure of aircraft fuselages is composed of skin, stringers and frames. The skin is reinforced longitudinally with stringers to reduce its thickness and so as to be competitive in weight, whereas the frames prevent general instability of the fuselage and may be subject to local loading. Within an aircraft fuselage we may encounter other structural elements, such as beams, which serve as framing for open sections of the fuselage or serve for supporting the loads introduced by the floor of the cabin of said aircraft.
Thus, the structure most used at present for a fuselage consists, on the one hand, of a skin with integrated, co-bonded or co-cured stringers and, on the other hand, of frames, and in their turn these frames can be floating or complete, being manufactured separately and being riveted subsequently to the skin of the fuselage.
The assembly of skin plus stringers can be manufactured in a single process (called one-shot), by which the skin, conical or cylindrical, is obtained together with the stringers in one piece, or alternatively said assembly of skin plus stringers can be manufactured separately in several panels (panelled solution) that are then joined together mechanically.
With regard to the frames, these can be, according to the prior art, floating or complete. In the case of the complete frames used at present, the manufacturing process is carried out in a large number of steps. The frames are manufactured separately, in several sections, and are joined mechanically to the skin, taking the form of said skin on resting on it. The problem posed by these known complete frames is that it is necessary to use complicated and expensive tooling to achieve the assembly tolerances required to enable said frames to be joined to the skin, taking into account the precise aerodynamic and structural requirements.
For the case of floating frames, the known manufacturing process also consists of several steps. The frames are manufactured separately but, apart from the sections that are required as a function of the panelling of the skin, the cross-section will consist of two different pieces: on the one hand, the floating frame as such and, on the other hand, the foot (piece called “babette” or “shear tie”) which is joined to the skin by means of rivets, and in its turn the floating frame itself is riveted to the aforesaid foot. With this solution, manufacturing of the floating frame is simplified, so that, as it is not necessary to copy the shape of the skin, the tooling used is simpler, and at the same time the problem of assembly tolerances is improved. However, this known solution of floating frames has the drawback of increasing the number of parts, and therefore the number of joints required.
In the two cases already known, complete frames and floating frames, different assembly stations are needed and a large quantity of fasteners (basically rivets), which involves weight penalties, high costs of production and assembly, and the need for greater logistic capacity.
That is why in recent years much effort has been devoted to achieving an ever increasing level of integration in the production of fuselages in composite material, so as to avoid the disadvantages of the aforementioned known solutions. The problem caused by this integration resides basically in creating sufficient pressure in all the elements during the process of combined curing.
As a result of these endeavours, there are several patents that describe methods of manufacture which, with the aid of special curing tooling, make it possible to integrate some of the typical structural elements, assembling the remaining elements in the next assembly stages. This is so for Patents WO2008/025860A1, WO2006001860A2 and US2006231682A1.
These patents focus on the tooling that makes it possible to manufacture the complete structure (skin in one piece, stringers and frame feet) in just one curing process.
Patent US2006231682A1 is based on laying-up the basic structural elements individually and, using suitable tooling, curing them together. The problem that arises from this document is that, as it is necessary to carry out the lay-up and forming of many parts, the costs of manufacture are very high, and furthermore, loads are transferred between the various basic structural elements through the bonded interface, and there may be problems of debonding in said joining zones.
The present invention aims to solve the disadvantages that have arisen previously.